The present invention relates to measuring or assessing the structural integrity of, the structural change in, or damage to, an aircraft component, and in particular to a load bearing metal element used as a safe-life component on an aircraft.
A safe-life component on an aircraft must be structurally safe throughout the entire working life of the component and in particular must not be allowed to develop cracks that could prejudice such structural safety, to yield to any non-negligible degree or to fail in any other way. (It will be understood that superficial micro-cracking on the surface of a safe-life component does not in itself constitute cracking that could prejudice structural safety. As such in the context of the present invention it will be understood that a component may be considered as being “crack-free” or “free from cracks”, despite the component having superficial micro-cracking, provided that such micro-cracking is of a nature that does not in itself prejudice structural safety.)
Safe-life components can be distinguished from “fail-safe” components, which are designed to be able to sustain damage or be allowed to develop cracks, to yield, or even fail, without presenting an unacceptable short-term safety risk. Fail-safe components may for example be able to sustain damage or partial failure without significantly affecting the ability of the component to perform its function, or if so significantly affected there are other components that are able to act as a back-up. For example, the failure of a safe-life component during service would be at risk of prejudicing safety to an unacceptable degree even in the short-term, whereas failure of a fail-safe component would be able to be tolerated until the next available opportunity for maintenance. Operators of aircraft typically have to conduct scheduled inspections of safe-life components at fixed intervals. Also, a safe-life component generally has to be replaced after a certain length of service so as to manage effectively the risk of failure of such a component in service. In view of the unacceptability of failure of a safe-life component during service, safe-life components are typically withdrawn far in advance of the possible maximum length of useable service and consequently the maximum length of service for safe-life components is typically conservatively short. Whilst such short lifetimes of safe-life components is wasteful, there is currently no means of effectively reducing the risk that a particular safe-life aircraft component will fail significantly in advance of the average lifetime. The present invention has been made in recognition that there has been a lack of effective means of assessing the structural integrity of such safe-life components once in use.
Various methods exist for measuring the structural integrity of a load bearing metal element including, for example, non-destructive testing methods (such as by means of X-ray radiography) and microscope analysis. Such techniques with their limited accuracy and other disadvantages, have limited application in relation to assessing and monitoring the structural integrity of a safe-life component. The present invention concerns the use of acoustic emission monitoring to assess the structural integrity of a safe-life aircraft component.
A method for detecting and monitoring fractures in a structure by monitoring acoustic emissions is disclosed in International Patent Application No. PCT/GB01/02213 (published under No. WO 01/94934) which is incorporated herein by reference in its entirety. The method of detecting and monitoring for damage in metal structures as described in WO 01/94934 relies on monitoring acoustic emissions caused by fractures or cracks. Such a method is therefore of no use when monitoring for damage in safe-life aircraft components, because it is a requirement that such components are free from cracks.
An experiment concerning the use of acoustic emissions to check the condition of a cylindrical specimen at high temperature is described in a paper entitled “Acoustic-Emission Study of Damage Accumulation During Alternating Load-Cycle Loading at Elevated Temperature” by N. G. Bychkov et al which is translated from “Problemy Prochnosti, No. 11, pp. 21-23, November 1983” of the “Central Institute of Aircraft Engine Design, Moscow and which is incorporated herein by reference in its entirety. The teaching of that paper relates primarily to monitoring of crack growth, and predicting imminent fracture in high temperature samples, in an experimental/laboratory setting utilizing a waveguide to carry acoustic emissions from the sample, which is housed in a furnace, to an acoustic emission transducer. As mentioned above, the present invention is concerned with structural health monitoring of safe-life components on an aircraft, such components being required to be free from cracks.